Gas turbine engines are widely applied machines for generating power or thrust. Most typically, they are employed on modern aircraft to provide the propulsion necessary for flight. They may also be used on board such aircraft for power generations or in an APU (Auxiliary Power Unit) capacity to provide for onboard heating, cooling, and ventilation, as well as the operational power and lighting systems onboard the aircraft within the cockpit, passenger cabin, and the like. They may also be used in land-based applications for generation of electrical power or mechanical horsepower in myriad vehicles and pieces of machinery.
In a typical gas turbine engine, three main sections are provided, namely, a compressor section, a combustion section, and a turbine section. Within the compressor section, ambient air is ingested, highly compressed, and directed through a downstream diffuser into the combustion section. Within the combustion section, the highly compressed air is mixed with fuel within an annular combustion chamber and burned at extremely high temperatures generating massive levels of heat energy. Moreover, as opposed to internal combustion engines, wherein the ignition of the fuel is intermittent every two or four strokes of the engine, ignition within a gas turbine engine is continuous, thereby increasing the high power levels attainable by the engine.
From the combustion section, the extremely hot combustion gases are directed to the turbine section downstream of the combustion chamber. As both the turbine section and the compressor section are mounted on the same shaft assembly, rotation of the turbine blades, upon contact with the rapidly expanding and hot combustion gases, causes the shaft to which they are mounted to rotate and in turn causes the compressor blades, also mounted to the shaft, to rotate and thus complete the engine cycle. The discharge of the rapidly expanding hot gases at high velocities from the turbine causes the engine to generate the aforementioned thrust needed for aircraft operation.
While effective, difficulties encountered in the design and operation of gas turbine engines result from the extreme temperatures to which the engine components, particularly the turbine blades, are exposed. Prior art devices have therefore devised schemes for directing cooling air into the turbine section of the gas turbine engine. One example is disclosed in U.S. Patent Application Publication No. U.S. 2002/0159880 A1 which teaches a plurality of orifices within the turbine section for directing cooling air axially through the gas turbine engine and then radially outwardly into the turbine section. Additionally, some cooling air is required to keep the rotor section and static parts of the turbine at acceptable temperatures. This cooling air, referred to as parasitic leakage air, is also injected radially outward into the turbine section.
While such devices are effective in directing cooling or parasitic leakage air into the turbine section, it is not without detrimental effect. More specifically, with such devices the cooling air or parasitic leakage is directed into the turbine section in a direction transverse to the gas flow path of the hot combustion gases moving through the turbine section. As the turbines are driven by this axial gas flow path of the combustion gases, any influx of air in a direction transverse to that flow necessarily detracts from the efficiency of the engine.
Accordingly, one of ordinary skill in the art can readily see that a need exists in the industry for a mechanism by which cooling air can be introduced into the turbine section of a gas turbine engine without significant detrimental effect to the operating efficiency of the engine.